Flight gas turbine with a first rotatable shaft

ABSTRACT

The present invention describes an aircraft gas turbine with a first rotatable shaft and a second shaft arranged coaxially thereto and which at least in an area close to the shaft end is non-rotatably connected to the first shaft. Recesses are provided in shaft areas at a distance from the connecting area between the shafts, said recesses overlapping one another at least in some areas when a defined twist of the first shaft relative to the second shaft is exceeded. The shafts are assigned at least one lever element designed rotatable and engaging in the recesses of the shafts in the radial direction in the area of an end of the shafts and swivelling about a pivot point. A second lever arm of the lever elements is, when the first lever arm is engaged in the recesses, operatively connectable to an electric sensor device, in the area of which a sensor signal equivalent to the defined twist of the first shaft can be generated.

This invention relates to an aircraft gas turbine with a first rotatable shaft in accordance with the type defined in more detail in the generic part of patent Claim 1.

A gas-turbine engine or an aircraft gas turbine with a first shaft arranged coaxially inside a second shaft is known from DE 33 22 430 A1. One of the shafts drivingly connects a first rotating assembly to a second rotating assembly. The second shaft is non-rotatably connected to the first shaft at least in an area close to the shaft end. In shaft areas at a distance from the connecting area between the shafts, recesses are provided which at least in some areas are overlapping one another when a twist of the first shaft relative to the second shaft, defined as a function of the torque applied in the area of the first shaft, is exceeded.

The shafts are, in the area of the recesses, assigned at least one lever element designed rotatable about a pivot point fixed relative to the shaft, which engages in the recesses of the shafts in the radial direction in the area of an end of a lever arm, when there is a corresponding, overlap of the recesses of the shafts, and swivels about the pivot point, with a radial distance between an end of a second lever arm of the lever element and the shafts increasing. This leads to one end of the second lever arm of the lever element being operatively connectable to an end of a further lever and likewise swivelling said lever. The swivelling of the end of the further lever has the effect that the lever is rotated with its arm about an axis. The rotary movement of the further lever in turn has the effect of shutting off the fuel supply of the aircraft gas turbine, so that the fuel supply in the direction of the combustion chamber of the aircraft gas turbine is ended. This measure is intended to prevent an increase in the speed of the low-pressure turbine of the aircraft gas turbine in the direction of an overspeeding range that might cause damage in the area of a gas-turbine engine.

One disadvantage is that the previously described safety system of a gas-turbine engine has an undesirably large number of mechanical parts, increasing the risk of the occurrence of failures. In addition, the known safety system for cutting off an engine during specific failures, e.g. when overspeeding states in rotating components occur, is also characterized by a high dead weight and a large installation space requirement, which however conflicts with aircraft design criteria.

Due to the purely mechanical design of the safety system in an aircraft gas turbine, the requirement placed on modern engine monitoring systems in the form of electronically verifiable or monitorable components using a so-called system check is not feasible. These verifications should wherever possible be performable every time the system is activated and also during operation of the system. Purely mechanical solutions are not suitable to do so. Exclusively electronically designed safety systems do permit electronic verification of functionality, but often have undesirably long triggering times, so that measures to prevent serious damage to engines might only be initiated when it is too late.

A device for an emergency shutdown of a gas turbine having electronically verifiable and mechanical components is disclosed by DE 197 27 296 A1 and includes a switchgear for actuating a control element that closes the fuel line when the axial displacement of the turbine shaft is not sufficient as a consequence of turbine overspeeding.

With this solution, however, there is the problem that the failure in the area of a gas turbine can only be detected when a shaft fracture occurs, so that damage causing high repair costs in the area of a gas turbine cannot be prevented by the device proposed here for an emergency shutdown of a gas turbine.

The object underlying the present invention is therefore to provide an aircraft gas turbine with a first rotatable shaft whose mode of operation is electronically verifiable and in which any impairment of the mode of operation can be detected within short operating times in a manner favourable in terms of installation space and at the same time with a low overall weight.

It is a particular object of the present invention to provide solution to the above problematics by an aircraft gas turbine having the features of patent Claim 1.

In the aircraft gas turbine according to the invention with a first rotatable shaft, via which a torque can be exchanged between at least two assemblies, and a second shaft arranged coaxially thereto and which at least in an area close to the shaft end is non-rotatably connected to the first shaft, recesses are provided in shaft areas at a distance from the connecting area between the shafts, said recesses overlapping one another at least in some areas when a twist of the first shaft relative to the second shaft, defined as a function of the torque applied in the area of the first shaft, is exceeded. The shafts are, in the area of the recesses, assigned at least one lever element designed rotatable about a pivot point fixed relative to the shaft and which engages in the recesses of the shafts in the radial direction in the area of an end of a lever arm, when there is a corresponding overlap of the recesses of the shafts, and swivels about the pivot point, with a radial distance between an end of a second lever arm of the lever element and the shafts increasing.

In accordance with the invention, the second lever arm is, when the first lever arm is engaged in the recesses of the shafts, operatively connectable to an electric sensor device at least at some times. In addition, a sensor signal equivalent to the defined twist of the first shaft can be generated in the area of the sensor device by making the operative connection.

The aircraft gas turbine in accordance with the invention is, in comparison with the purely mechanical safety system known from the state of the art, designed with a smaller number of components and therefore has a smaller installation space requirement while being characterized by a lower component weight. In addition, a function of the safety system of the aircraft gas turbine is electronically verifiable to the required extent using the electric sensor device, which is not possible with the known and purely mechanical design.

Compared with the known device for an emergency shutdown of a gas turbine that has both mechanical and electronic components, failures in the aircraft gas turbine in accordance with the invention can be detected before the occurrence of massive damage in the area of the aircraft gas turbine with the desired short triggering times, and damage in the area of the aircraft gas turbine during impermissible operating states of an aircraft gas turbine can be prevented by timely initiation of suitable counter-measures to the required extent.

With embodiments of the aircraft gas turbine in accordance with the invention that are favourable in terms of both installation space and costs, and that can be operated at low expenditure, the sensor device includes at least one electric switch or one electric sensor with cabling that can be brought into contact with the second lever arm.

If the electric sensor is designed as a wire sensor and if the second lever arm has a blade area operatively connectable to a signal wire, by means of which said signal wire can be severed, the sensor signal equivalent to the defined twist of the first shaft can be generated with low design expenditure and at the same time electronic verification of the mode of operation of the safety system of the aircraft gas turbine can be achieved in a simple manner.

the electric sensor is designed as a force sensor, using which the sensor signal equivalent to the defined twist of the first shaft can be generated upon contact between the lever arm and the force sensor, operating states causing damage in the area of an aircraft gas turbine can be prevented with low expenditure due to short triggering times, and in addition the mode of operation of such a safety system of an aircraft gas turbine is electronically verifiable to the required extent.

In a simply designed embodiment of the aircraft gas turbine in accordance with the invention, the sensor device is linked to a fuel shutoff valve, by means of which a fuel supply can be interrupted when the sensor signal equivalent to the defined twist of the first shaft is given. As a result, overspeeding states in rotating parts of the aircraft gas turbine can be prevented in a simple manner when a failure is detected, since a drive unit of the aircraft gas turbine can be shut down within short operating times.

Further advantages and advantageous developments of the present invention become apparent from the patent Claims and the exemplary embodiment described in principle with reference to the drawing. Here,

FIG. 1 shows a highly schematized longitudinal sectional view of an aircraft gas turbine,

FIG. 2 shows an enlarged view of an area II of he aircraft gas turbine, said area II being indicated in more detail in FIG. 1, and

FIG. 3 shows an enlarged representation of an area III of the aircraft gas turbine from a view indicated in more detail in FIG. 2.

FIG. 1 shows an aircraft gas turbine 1 or a jet engine of an aircraft respectively, having a bypass duct 2 and an inlet area 3. A fan 4 follows in a manner known per se downstream of the inlet area 3.

Again downstream of the fan 4, the fluid flow inside the jet engine 1 splits into a bypass flow and a core flow, with the bypass flow flowing through the bypass duct 2 and the core flow into an engine core 5. The engine core 5 is designed with a compressor device 6, a burner 7, a low-pressure turbine 8 intended for driving the fan 4 and a high-pressure turbine 9 intended for driving the compressor device 6. In addition, FIG. 1 shows a schematically illustrated accessory gearbox 10, which is arranged substantially in the area of an intermediate casing 11 of the jet engine 1 and which can, depending on the application in question, also be provided in another suitable area of the jet engine 1. The intermediate casing 11 is located in the radial direction of the jet engine 1 in an area between the engine core 5 and the bypass duct 2.

The accessory gearbox 10 is driven by a drive shaft 13 interacting with an engine shaft 12, arranged substantially parallel to the engine shaft 12 and in the present invention operatively connectable to the engine shaft 12 via an auxiliary shaft 14. The auxiliary shaft 14 is connected via a bevel gearing 15 to the engine shaft 12, and interacts with a high-pressure shaft which rotates during operation of the jet engine 1 at a higher speed than a low-pressure shaft 16 coaxially arranged thereto and connected to the fan 4. The fan 4 is non-rotatably connected to the low-pressure turbine 8 via the low-pressure shaft 16, so that a torque can be transmitted via the low-pressure shaft 16 between the fan 4 and the low-pressure turbine 8. Coaxially to the low-pressure shaft 16, a further, second shaft 17 is provided inside the low-pressure shaft 16 and non-rotatably connected to said low-pressure shaft 16 in the area of the fan 4.

FIG. 2 shows an enlarged representation of an area II indicated in more detail in FIG. 1, including the low-pressure turbine 8 and ends of the low-pressure shaft 16 and the further shaft 17 facing away from the fan 4. In addition, FIG. 3 shows an area III indicated in more detail in FIG. 2 and including the ends of the low-pressure shaft 16 and the further shaft 17.

As can be seen from the representations in accordance with FIG. 2 and FIG. 3, recesses 16A, 17A are provided in areas of the low-pressure shaft 16 and the further shaft 17 at a distance from the connecting area between the low-pressure shaft 16 and the further shaft 17 or in the area of their shaft ends, said recesses overlapping one another at least in some areas when a twist of the low-pressure shaft 16 relative to the further shaft 17, defined as a function of the torque applied in the area of the low-pressure shaft, is exceeded. The low-pressure shaft 16 and the further shaft 17 are, in the area of the recesses 16A and 17A, assigned here two lever elements 20, 21, each designed rotatable about a pivot point 18, 19 fixed relative to the shaft, which engage in the recesses 16A and 17A of the low-pressure shaft 16 and the further shaft 17 in the radial direction in the area of an end 20A or 21A respectively, of a lever arm 20B, 21B, when there is a corresponding overlap of the recesses 16A, 17A of the low-pressure shaft 16 and the further shaft 17, thereby swivelling about the pivot points 18 and 19.

During swivelling of the lever elements 20 and 21 about the pivot points 18 and 19, a radial distance increases between an end 20C/21C of a second lever arm 20D/21D respectively. The previously described rotary movement of the lever elements 20 and 21 is triggered in the present invention by spring elements 22 and 23 and by centrifugal forces acting on the lever elements 20 and 21 during operation, if the rotary movement has been released by the previously described overlapping of the recesses 16A and 17A.

FIG. 3 shows the lever element 21 in an operating state, where the end 21A of the lever arm 21B only engages in the recess 16A of the low-pressure shaft 16, since the torque-dependent twist of the low-pressure shaft 16 relative to the further shaft 17 is smaller than the defined twist, where the recesses 16A and 17B are in an overlapping operating state necessary for passage of the end 21A of the lever element 21 through both the recesses 16A and the recesses 17A. By contrast, the lever element 20 is, due to passage of the end 20A through both recesses 16A and 17A of the shafts 16 and 17, swivelled about the pivot point 18, and the second end 20C of the lever element 20 is at a greater distance from the surface of the low-pressure shaft 16 in the radial direction than the second end 21C of the lever element 21.

The second lever arms 20D and 21D of the lever elements 20 and 21 are, when the first lever arms 20B and 21B are engaged in the recesses 16A and 17A of the low-pressure shaft 16 and the further shaft 17, operatively connectable to an electric sensor device 24 at least at some times, where in the area of the sensor device 24 a sensor signal equivalent to the defined twist of the low-pressure shaft 16 can be generated by making the operative connection.

The sensor device 24 is designed in the present invention as a so-called wire sensor having a signal wire 25. Due to the rotation of the low-pressure shaft 16 and the further shaft 17, the lever element 20 or 21 severs the signal wire 25 with a blade area 20E or 21E respectively, which first reaches the signal wire 25 with its swivelled end 20C or 21C of the second lever arm 20D or 21D respectively. Hence a signal can be triggered by the sensor device 24, by means of which a fuel shutoff valve of the aircraft gas turbine 1 is actuated and a fuel supply to the aircraft gas turbine is interrupted.

Depending on the application in question, it is also possible to arrange several sensor devices 24 spread over the circumference of the low-pressure turbine 16 in order to remain capable of detecting a failure in the aircraft gas turbine 1 when one of the sensor devices 24 is lost, for example due to another failure.

The aircraft gas turbine in accordance with the invention combines the advantages of purely mechanical and purely electronic safety systems for detecting failures and for triggering counter-measures, by means of which undesirable damage to aircraft engines, for example caused by overspeeding states of rotating parts such as turbine rotors and the like, can be prevented in simple manner within short operating times. To do so, a conventional reference angle tube in conjunction with a sensor device preferably designed as an electric switch or sensor with appropriate cabling is used for determination of the torsion angle or a twist of a torque-transmitting shaft of an aircraft gas turbine or of a reference angle position of the rotors. When the torsion angle of the torque-transmitting shaft in the area of an end section of a reference tube or of the further shaft is exceeded, recesses or slots of the torque-transmitting shaft and of the reference tube are superimposed, so that one or more levers can engage in the manner previously described. An opposite lever arm end of a lever element or of a tilt lever is moved radially outwards by centrifugal forces acting on it and severs a signal wire, preferably by means of a blade. A signal can be triggered by the severing of the signal wire and can be used to actuate a fuel shutoff valve.

LIST OF REFERENCE NUMERALS

-   1 Aircraft gas turbine, jet engine -   2 Bypass duct -   3 Inlet area -   4 Fan -   5 Engine core -   6 Compressor device -   7 Burner -   8 Low-pressure turbine -   9 High-pressure turbine -   10 Accessory gearbox -   11 Intermediate casing -   12 Engine shaft -   13 Drive shaft -   14 Auxiliary shaft -   15 Bevel gearing -   16 Low-pressure shaft -   16A Recess of low-pressure shaft -   17 Further shaft -   17A Recess of further shaft -   18, 19 Pivot point fixed relative to the shaft -   20 Lever element -   20A End of first lever arm -   20B First lever arm -   20C End of second lever arm -   20D Second lever arm -   20E Blade area of second lever arm -   21 Lever element -   21A End of first lever arm -   21B First lever arm -   21C End of second lever arm -   21D Second lever arm -   21E Blade area of second lever arm -   22, 23 Spring element -   24 Sensor device -   25 Signal wire 

1. Aircraft gas turbine with a first rotatable shaft, via which a torque can be exchanged between at least two assemblies, and a second shaft arranged coaxially thereto and which at least in an area close to the shaft end is non-rotatably connected to the first shaft, where recesses are provided in shaft areas at a distance from the connecting area between the shafts, said recesses overlapping one another at least in some areas when a twist of the first shaft relative to the second shaft, defined as a function of the torque applied in the area of the first shaft, is exceeded, where the shafts in the area of the recesses, are assigned at least one lever element designed rotatable about a pivot point fixed relative to the shaft and which engages in the recesses of the shafts in the radial direction in the area of an end of a lever arm, when there is a corresponding overlap of the recesses of the shafts, and swivels about the pivot point, with a radial distance between an end of a second lever arm of the lever element and the shafts increasing, characterized in that the second lever arm is, when the first lever arm is engaged in the recesses of the shafts, operatively connectable to an electric sensor device at least at some times, and a sensor signal equivalent to the defined twist of the first shaft can be generated in the area of the sensor device by making the operative connection.
 2. Aircraft gas turbine in accordance with claim 1, wherein the sensor device includes at least one electric switch or one electric sensor with cabling that can be brought into contact with the second lever arm.
 3. Aircraft gas turbine in accordance with claim 2, wherein the electric sensor is designed as a wire sensor and the second lever arm has a blade area operatively connectable to a signal wire, by means of which said signal wire can be severed.
 4. Aircraft gas turbine in accordance with claim 1, wherein the electric sensor is designed as a force sensor, using which the sensor signal can be generated upon contact between the second lever arm and the force sensor.
 5. Aircraft gas turbine in accordance with claim 1, wherein the sensor device is linked to a fuel shutoff valve, by means of which a fuel supply can be interrupted when the sensor signal equivalent to the defined twist of the first shaft is given. 